With the increasing use of composite materials in the manufacture of aircraft components, a number of problems have arisen in connection with the application to such materials of design principles developed for conventional aluminum construction. Such design principles involve the use of numerous mechanical fasteners and the inclusion of multiple details. Advanced composite materials do not readily lend themselves physically or economically to extensive use of mechanical fasteners and highly detailed designs. The inclusion of multiple details in a composite material component increases the cost of manufacture to unacceptably high levels and makes the application of automated procedures difficult, if not impossible. Mechanical fasteners are not physically well suited for composite material components, and their use with such components tends to detract from the advantages of the unidirectional properties of advanced composite fibers. These problems have created a need for developing new techniques for economically producing composite material components, especially such components that are required to be load resisting.
In recent years there have been a number of proposals relating to the structure and manufacture of composite material components. Some of these proposals involve the use of thermosetting adhesives instead of mechanical fasteners. Manufacturing processes that employ thermosetting adhesives generally require the separate forming and curing of the components and a further heating process to set the adhesives. Such multiple step processes have the disadvantages of being time consuming and expensive to carry out. Other proposals involve the separate forming of elements of a component and the curing together of the elements to form the desired structure. A serious problem encountered in connection with known composite material components embodying both approaches--bonding by adhesives and bonding by curing--has been the tendency for one part of the component to be peeled away from an adjacent part of the component when the component is subjected to stress forces. Recent proposals for overcoming such peel tendencies have the serious drawback of being quite complicated and expensive to carry out.
An aircraft component that is at least partially formed from composite materials and/or a process for making such a component is disclosed in each of the following U.S. Pat. Nos.:
3,096,958, granted July 9, 1963, to R. D. Koontz; 3,519,228, granted July 7, 1970, to L. J. Windecker; 3,645,829, granted Feb. 29, 1972, to Palfreyman et al; 3,768,760, granted Oct. 30, 1973, to L. C. Jensen; 3,780,969, granted Dec. 25, 1973, to Nussbaum et al; 3,902,944, granted Sept. 2, 1975, to Ashton et al; 3,962,506, granted June 8, 1976, to E. O. Dunahoo; 3,995,080, granted Nov. 30, 1976, to Cogburn et al; 3,995,081, granted Nov. 30, 1976, to Fant et al; and 4,310,132, granted Jan. 12, 1982, to Robinson et al.
Koontz discloses a sheet material with a plurality of generally planar ribs spaced along one of its surfaces. The sheet material is made from a continuous filament composite material, and the ribs are made from a composite material with discontinuous, randomly-disposed filaments. The ribs are positioned on the sheet material during a rib molding process, and the sheet material and the ribs are cured together to bond them together.
Windecker discloses an airfoil structure that includes a foamed core, a composite skin, and spaced planar spars made from laminates of reinforced plastic and metal. These three elements are "adhered" together.
Cogburn et al and Fant et al each disclose a composite material plicated structural beam with a fairly complex design that is described as being peel resistant. Each of the elements of the beam is formed separately, and then the elements are assembled and cured together to form the beam. The use of a destructible mandrel in the curing process is described.
Robinson et al disclose a fuselage structure having a number of spaced stringers and reinforcing members. The composite skin of the structure has layers oriented at plus and minus 45.degree. degrees to the fuselage longitudinal axis, and the stringers and reinforcers have layers oriented at 0.degree. to such axis. The stringers and reinforcing members are secured to the skin by means of mechanical fasteners, brazing, or compacting.
Nussbaum et al disclose a wing case for airfoils having a laminated skin and spaced apart metal spars and ribs. The spars and ribs are secured to the skin by mechanical means. The layers of the laminated skin are oriented at 0.degree., 90.degree., and plus and minus 45.degree. to the span of the wing.
Jensen discloses a laminated composite skin for covering a structural component. The layers of the skin are oriented at a plurality of angles, and some of the layers have fibers oriented in more than one direction.
Palfreyman et al disclose apparatus for producing composite material by a filament winding process. The manufacture of an airfoil shape is described. Ashton et al disclose a filament winding process for forming a flexible sheath which is then removed from the mandrel and molded into a desired noncircular shape.
Dunahoo discloses a composite material airfoil with a multichambered cellular structure. The manufacture of the airfoil involves multiple winding and curing processes.
The above patents and the prior art that is discussed and/or cited therein should be studied for the purpose of putting the present invention into proper perspective relative to the prior art.